The modern gas turbine engine flow path undergoes a dramatic series of axial temperature changes from fan inlet to low turbine discharge. Because the flow path is a near closed system, isolated from the components inboard and outboard, the blades, vanes, and disk rims contained within the system reach operating temperatures up to 2500° F. (surface temp.) in the gas path, and 2500-3200° F. rim temperatures in the high turbine. Conversely, the disk webs and bores which are inboard of the flow path are located in cavities isolated by rotating labyrinth seals and pressure balanced to prevent ingestion of hot gasses into the engine core.
Secondary flow systems, precisely designed, keep disk bores and drive shafts at lower temperatures to limit blade radial growth and maximize material structural properties. Maintaining and controlling thermal gradients between disk rims and bores controls transient disk/blade growth and minimizes tip rubbing and resultant aerodynamic performance loss.
The rim-to-bore thermal gradients represent an additional stress component in conjunction with the rotational stress. In many cases the thermal stress contribution can exceed the mechanical stress. There is also phase lag between mechanical and thermal stresses. The more isolated and massive the disk bore, as in the high turbine, the greater the phase lag becomes. The phase problem is also affected by operability since power manipulation, to a certain extent, determines the severity of the gradient. This creates a situation where different operators may experience inconsistent rotating component life strictly based on how they operate their engines.
Structural integrity personnel have the challenge of trying to establish a life limit for components with simultaneously changing elastic/inelastic strains and material behavior. Data derived from material coupon testing is of limited value because there is no current material testing method (outside of an actual jet engine) that induces a radial thermal gradient in synchronization with realistic in-service rotational speeds. In other words, there is no method that simulates a thermal and mechanical fatigue (TMF) load cycle.
Given the extensive testing used to generate durability and life prediction models, and the cost of full-up engine tests ($6M+each), no cost-effective way to conduct representative TMF testing is presently available. Current component design methods are also very costly, often requiring several iterations to achieve a “best effort” solution, resulting in increases of many millions of dollars and months of schedule to an engine program.
Additionally, TMF related failures in engine hot sections are a critical factor driving class A mishaps, and low Total Accumulated Cycles (TAC) useful life, Mean Time Between Failure (MTBF), and Mean Time Between Maintenance (MTBM). Advances in design that come from better testing have the potential to produce components that are more resistant to the effects of TMF, helping more advanced engines reach 4300 TAC life for the hot section, and help legacy engines extend hot section life to 5000 or perhaps 6000 TACs. Such improvements in component durability could save the USAF hundreds of millions of dollars in future maintenance cost avoidance. Combined DoD and commercial market savings might reach billions of dollars.
Analytical based life prediction systems have fallen short in assessing component life with the result of imposing financial burdens on operators for replacement parts. While analysis tools exist that can model engine components and TMF loading, the results of such model analyses may not adequately predict the response of objects and materials subjected to the environment typically encountered in the gas turbine engine. For example, estimates or assumptions made in constructing such a model may have an exaggerated or underrepresented impact on the test results. Correlation between physical test results and predicted results can sometimes vary widely, often leading designers to specify conservative estimates for predicted lifetimes of engine components.
Commercial and military hot section components are typically retired after a specified number of cycles, or TACs, for which a given hot section component is rated. The particular number of cycles or TACs for which a component is rated is based in part on the particular component. It is commonly believed that there is additional useful life in components retired for reaching the specified number of cycles of TACs. For example, engineering conservatism due to lack of knowledge regarding material behavior under severe TMF conditions typically results in shorter prescribed lives for hot section rotational components. Programs such as ERLE (Engine Rotor Life Extension) seek to validate, through better testing methods, that current component designs do indeed have longer useful lives. It is predicted that significant cost savings per year could be realized by not retiring components too early.
To date, spin pit based TMF testing has been somewhat impractical due to the lack of high heat flux ovens capable of duplicating engine thermal excursions. Prior tests have been conducted with cyclic speeds but constant gradients. The constant gradients typically have been created with induction or resistive element heaters. In the case of turbine hot section components, rim temperatures can reach 2500° F. or higher.